1. Field of the Invention
The present invention relates to a thermal balancing system, and in particular to a thermal balancing system and method for terrestrial thermal vacuum testing of spacecraft to simulate on orbit conditions by the replication of the heat transport functions of heat pipes which are dysfunctional due to gravitational forces. The present invention also relates to a system and method for preheating a spacecraft prior to terrestrial thermal vacuum testing, and a method for design of a cooling plate utilized by the above systems.
2. Background of Related Art
Unmanned space vehicles, in particular satellites, are now used for everything from telecommunications to weather reporting to missile defense. With this growing demand has come increased size, complexity, and capability resulting in greater cost. As a result, ever increasing amounts of terrestrial, ground-based, testing are required to ensure that such spacecraft will function properly throughout their effective life.
In its basic elements, a spacecraft can be broken up into six functional groups or subsystems: the spacecraft structure; the power system; the attitude control system; the telemetry, tracking, command and communications system; the propulsion system; and the thermal control system. The thermal control system is an important element as its purpose is to achieve thermal balance of the spacecraft, permitting proper performance of all subsystems during the spacecraft's proposed operational modes. As one might expect, through the various phases of a spacecraft's mission there are variations in heat generation and dissipation due to internal component operation and surface fluxes caused primarily by solar heating and eclipse. As each on board system has a thermal operating range, it is the responsibility of the thermal control system to maintain the temperature of subsystems within their best operating range.
During the initial design phase of a spacecraft, analytical thermal system models are developed. The models not only predict a thermal system balance temperature but also define heat transfer paths for which heat dissipation devices can be strategically placed to discard the waste heat. While these models are adequate for design, the predictions must be verified by actual system testing.
As there is no or negligible air at spacecraft orbital altitudes, the primary heat transfer mechanisms are conduction and radiation. Conductive heat transfer is achieved through material and surface-to-surface contact, while radiative heat transfer is primarily achieved through the use of radiator panels which radiate the heat into deep space. Excess heat generated by internal components or solar heating is conducted from the source to the radiator surfaces, which then rejects the heat to space. While there are a variety of convective heat transfer methods to move the excess heat from the source to the radiative panels, one preferred method is a heat pipe. The advantage of heat pipes is that they typically exhibit thermal conductivity greatly in excess of most thermally conductive metals and transport thermal energy at efficiencies greater then 90%. A heat pipe is a closed loop system which contains an evaporator section, a condenser section, and a working fluid. Should external geometric requirements make it necessary, an adiabatic transport section can be included to further separate the evaporator and condenser portions. Internal to the heat pipe are typically a series of axial grooves which provide a transport path for condensed fluid to flow from the condenser section to the evaporator section.
The heat pipe is mounted such that the evaporator section is attached to a heat source whose temperature is to be maintained within a predetermined range. The condenser section is generally attached to a heat sink such as a space heat radiator for radiation of the heat to space. It is not uncommon for a spacecraft to have several heat pipes, or banks of heat pipes, located throughout the vehicle and oriented in a variety of positions depending upon heat sources and heat sinks.
In operation, the heat generated by a heat source is absorbed by the working fluid in the evaporator section of the heat pipe, vaporizing the working fluid. The vaporization causes an increase vapor pressure difference resulting in the vaporized fluid traveling, inside the heat pipe, from the evaporator section to the condenser section. Since the condenser section is attached to a heat sink, the vaporized fluid reaching the condenser section is cooled, causing it to condense and release the latent heat of vaporization. The condensed working fluid is transferred back to the evaporator section by capillary action, where it is again evaporated, absorbing the heat from the evaporator section.
In a substantially gravity-free space environment, the transfer of the working fluid over the length of the heat pipe is not a problem. However, on Earth gravity may inhibit the return of the working fluid, especially if the condenser section is located below the evaporator section. This is because capillary action of the working fluid in a one-G environment is generally limited to a vertical rise of about 13 millimeters or less. As a result, the functionality of a heat pipe in a terrestrial environment is directly affected by its orientation.
Terrestrial thermal system testing of spacecraft, otherwise known as thermal vacuum testing, is typically done in a large vacuum chamber capable of housing an entire satellite at vacuum levels lower than of 1.times.10.sup.-5 torr. As a result of the limited capillary rise of a working fluid in a one-G environment, thermal vacuum testing can be quite difficult and give incomplete results. This is due to the fact that often one or more of the heat pipes are oriented in such a manner that they are functionally inoperable because of the gravitational force. While data can be acquired by those systems that are functional, such testing cannot develop a complete thermal balance for the entire vehicle. As a result, either multiple tests run must be conducted with the spacecraft manually repositioned to different orientations, or the spacecraft must be mechanically rotated during test. While these approaches provide additional data with respect to the functionality of the thermal control system, such testing can still not provide a complete thermal balance since not all systems are functional at the same time. In addition to the inability to provide a thermal balance, such alternate approaches are costly both in material and man-hours. Each manual repositioning of the satellite requires the test chamber to be vented to the atmosphere, the satellite rotated, a full pump-down cycle completed, and the spacecraft thermally stabilized. It can take from 12 to 48 hours between successive tests. If the satellite is rotated within the chamber during test, special tooling must be designed and manufactured. This tooling must not only be capable of supporting and rotating the satellite during test, but also withstanding the thermal and vacuum environment.
There is thus a need for method that would permit full thermal system testing of spacecraft in a single test cycle and provide a system thermal balance. Additionally, such a system should be able to provide a full system balance within a margin of error of 10% or less, preferably 5% or less.